1. Field of the Invention
This invention relates to gas turbine engine cracks and, more particularly, to techniques to divert cracks from sensitive areas and also so that the cracks propagate in a more benign mode than would otherwise occur.
2. Description of Related Art
Gas turbine engine stationary structures and components as contrasted to gas turbine engines rotors are subject to cracking due to thermal, pressure and vibratory or flight loads which subject the components to fatigue failure and reduced life. Stationary turbine vane airfoils, typically made from Titanium materials, are subject to stress due to various excitations of the airfoil in bending and torsional flexure modes. Airfoil failure can result from these responses of the airfoil to combinations of chordwise and spanwise bending modes resulting from gaseous flow impingement, rotor blades passing by stator vanes and flutter. These excitation responses are characterized by nodal patterns on the airfoil resulting in panel sections of the airfoil, or other stationary plate type components, vibrating at resonant frequencies. Cracks form and propagate along nodal lines of the nodal patterns. Nodal lines are linear or curvilinear lines about which portions of the vane airfoil or plate bend and are therefore subject to maximum stress levels. Different nodal lines, modes, and modes of failure are characterized by their order of resonance, i.e. 1st, 2nd, 3rd, etc..
Different modes of failure also can cause different degrees of damage some less severe or more benign than others. For example, a variable angle vane has a trunioned vane and a crack which progresses from a crack initiation point due to an 8th order failure mode sometimes shifts (depending on vane thickness variations) to a 9th order failure mode. The 8th order failure mode results in the release of a lower corner of the vane airfoil causing minor damage to downstream compressor blades and engine components and experience shows that the engine continues to run. The 9th order failure mode results in release of a lower trunnion portion which is significantly larger than the portion released by an 8th order failure mode and has been found to result in considerably more downstream damage and may even result in an engine failure or shutdown. Therefore, it is also highly desirable to design and construct gas turbine engine static components which are susceptible to more than one resonant mode of failure such that should a failure occur it is the one which causes less damage or is more benign than the other. The present invention is directed towards this end and provides a gas turbine engine static component and in particular a trunioned variable stator vane airfoil with at least one region of deep compressive residual stresses imparted by laser shock peening disposed in an area between two nodal lines of two respective resonance failure modes.
Another area of concern as regards to the present invention is cracks that originate at features of casings and then propagate into sensitive or area that are more vulnerable to fatigue induced failure. Combustor cases have annular shells or panels that axially extend between integrally formed annular flanges which are bolted to adjacent flanges of adjacent casings. Certain compressor casings such as those found on a General Electric T39 have horizontally split cases with horizontal axially extending flanges which terminate at annular flanges which are bolted to the annular flanges of the adjacent combustor casing. The points at which the horizontal axially extending flanges terminate at annular flanges is a feature which causes high localized stress in the annular flange of the combustor casing which can cause a crack to initiate in the flange and propagate to the point of fatigue failure in the shell of the combustor casing. The shells are relatively thin as compared to the flanges and are subject to fatigue failure because of thermal and pressure loading. A fuel nozzle boss on the combustor casing shell is another feature that is capable of initiating a crack that can propagate into areas of the shell that are more susceptible to fatigue failure, such as the hotter running areas that are axially downstream of the fuel nozzles. Therefore, it is highly desirable to design and construct gas turbine engine static components such as combustor casing which are susceptible to such cracks with means to divert cracks that divert the cracks which are initiated from certain features attached to the shell and prevent them from propagating into areas of the shell that are more susceptible to fatigue failure. The present invention is directed towards this end and provides a gas turbine engine static combustor shell with at least one region of deep compressive residual stresses imparted by laser shock peening disposed in an area between such crack initiating features and the shell of the component casing.
The region of deep compressive residual stresses imparted by laser shock peening of the present invention is not to be confused with a surface layer zone of a workpiece that contains locally bounded compressive residual stresses that are induced by a hardening operation using a laser beam to locally heat and, thereby, harden the workpiece such as that which is disclosed in U.S. Pat. No. 5,235,838, entitled "Method and Apparatus for Truing or Straightening Out of True Work Pieces". The present invention uses multiple radiation pulses from high power pulsed lasers to produce shock waves on the surface of a workpiece similar to methods disclosed in U.S. Pat. No. 3,850,698, entitled "Altering Material Properties"; U.S. Pat. No. 4,401,477, entitled "Laser Shock Processing"; and U.S. Pat. No. 5,131,957, entitled "Material Properties". Laser peening as understood in the art and as used herein means utilizing a laser beam from a laser beam source to produce a strong localized compressive force on a portion of a surface. Laser peening has been utilized to create a compressively stressed protection layer at the outer surface of a workpiece which is known to considerably increase the resistance of the workpiece to fatigue failure as disclosed in U.S. Pat. No. 4,937,421, entitled "Laser Peening System and Method". However, the prior art does not disclose fan blade leading edges of the type claimed by the present patent nor the methods of how to produce them. It is to this end that the present invention is directed.